A solar array, as defined herein, pertains to a structure which is stowable in a small volume for shipment and launch, and that is deployable when in space to expose a large surface area of photovoltaic collectors (solar cells) to the sun, and that is attached to certain spacecraft vehicles, to provide power for spacecraft operations. Solar arrays typically consist of an underlying structure for deployment of a substantial number of individual photovoltaic solar cells from the body of a spacecraft. Once fully deployed, it is desirable for the solar array structure to provide a lightweight, stiff, strong, stable and flat platform for the solar cells to allow uniform exposure to the sun and minimize on-orbit spacecraft attitude control disturbance loads. Solar arrays are typically stowed such that they are constrained in a small envelope prior to and during launch of the spacecraft and then are deployed to their fully extended configuration, exposing the maximum area of solar cells once the spacecraft has reached its position in outer space. It is desirable to minimize the volume of the stowed package while at the same time maximizing the available solar cell area that can be packaged when stowed, and subsequently deployed to allow for maximum power production for the spacecraft.
In certain spacecraft and other space-based solar power applications, an extremely large area of photovoltaics needs to be deployed to provide very high levels of power. While the largest solar arrays deployed from a typical large communication satellite provide 15 to 25 kilowatts (kW) at beginning of life, an very-high power solar array may be required to provide from 100 kW to greater than 1000 kW (1 megawatt) of power. An example of missions requiring such large solar array power are solar electric propulsion (SEP) missions that utilize high power levels to run an ion-thruster engine for space propulsion.
FIG. 1 shows a typical spacecraft (101) that uses a high-power solar array (102) for power production, with the high-power solar array (102) in the deployed configuration.
As the deployed size of solar arrays is scaled up to a degree required for ultra-high power applications, the need to package very efficiently from the deployed state into a sufficiently small stowed volume becomes critical and enabling to allow the significant solar cell area to be launched into space. This is because the limiting constraint is the available size launch vehicle fairing volume that carries the stowed high-power solar array, spacecraft bus and all associated mission hardware in a single earth-to-space launch. FIG. 2 shows a typical spacecraft (101) that utilizes a high-power solar array (102) in the configuration where it is stowed (packaged for launch, 201) and within a typical launch vehicle shroud (202), showing the stowed volume within the shroud available for the stowed high-power solar array (2.03).
In many prior art applications of solar arrays, the structure consists of flat rigid panel substrates that are configured for stowage by means such as hinges which will permit the panels to be folded against each other to minimize the dimensions of the array in the stowed configuration. The stowed packaging efficiency (defined as the ability to fill up or utilize the available volume for the stowed solar array inside the launch vehicle volume) of the typical folded-up rigid panel solar arrays becomes poor as the array is scaled up to high-power levels. The stowed packaging of a very large high-power rigid panel solar array involves the use of many more moving mechanical items such as hinges and latches; and actuating mechanisms such as springs, cables and pulleys. The much greater number of mechanical components required for a high-power rigid panel application reduces deployment reliability, and increases system weight and cost. The rigid panels themselves add significant undesirable weight when scaled up to very large sizes or when their numbers increase to meet deployed area requirements.
Other key considerations when scaling up solar arrays to very high power levels is the minimization of weight, and the maximization of deployed stiffness and strength. Because of its much larger size, the proportion of launch mass of a high-power solar array is much higher portion of the overall spacecraft mass, limiting options for launch vehicles to those capable of carrying sufficient mass to orbit. A lower mass, smaller packaged alternative to the state-of-the-art solar arrays allows a larger selection of available launch vehicles to be utilized for very high-power missions. It is also desirable to maximize the deployed natural frequency (stiffness) and strength (against deployed accelerations) of a solar array. Low mass and high stiffness/strength when deployed results in sufficiently low deployed mass moments of inertia and high deployed frequencies which enable standard “passive” methods of spacecraft attitude control systems and more accurate sun-pointing of the large-area structures to be implemented. As the size of the solar cell deployed area and the solar array supporting structure increase, the stiffness of the solar cell array decreases and, as a result, the vibration frequency decreases and disturbance deflections increase. The ability of the spacecraft attitude control system to orient the spacecraft may be impaired if the deflections due to low-frequency solar array movement are excessive.
In order to allow for the added reduction in a deployable solar arrays weight and stowed volume required for very high power applications, the solar cell mounting can be configured using a flexible substrate, or blanket. Various flexible solar cell blanket substrates have been used, such as those fabricated from a fiberglass mesh or thin polymeric sheet upon which are bonded the numerous crystalline solar cells. Flexible-blanket solar arrays for use on spacecraft have typically been limited to crystalline solar cell arrays packaged in a long roll or pleated stack that is deployed using a separate deployment boom or hub structure requiring external motor power for deployment motive force. These flexible array deployment structures have consisted of very complex mechanical systems such as coilable or articulated truss booms, or radially oriented spars that rotate about a central hub, which can add undesired parts, complexity, weight and cost to implement. Examples of prior art flexible blanket arrays are shown in the following U.S. patents: Harvey et al U.S. Pat. No. 5,296,044; Stribling et al U.S. Pat. No. 6,983,914; and Hanak et al U.S. Pat. No. 4,636,579.
Because of the extreme size (and corresponding weight) required of a typical very high power solar array, both during deployment and when fully deployed, the ability to verify the solar array mechanical and electrical function on earth, under one earth gravity (1-G) through functional testing becomes a primary consideration. The system required for 1-G off-loading support of the high-power solar array to simulate zero-gravity during deployment for all critical elements of the huge array structure can rival the array itself in complexity and cost, and in some cases it is physically impossible to off-load/support such a large, gossamer structure during deployment under 1-G to adequately simulate a deployment in space. Additionally the photovoltaic elements and electrical performance of the large solar cell-populated area needs to be validated and verified throughout the ground test and integration phases prior to launch. A key consideration in the viability of any high solar array power design is the ability to test-validate its performance on the ground.
A review of the prior art of large solar arrays shows that significant efforts have been made to reduce the weight and increase the deployment reliability of rigid panel and flexible blanket-type solar arrays for a given set of deployed stiffness and strength requirements. Although these prior large solar array design solutions have resulted in solar arrays tending to involve difficult and time consuming manufacturing, higher complexity and higher cost, most of these prior applications were designed for power applications below 30 kW. Most of the prior-art solutions therefore do not scale-up to the degree needed for very high power (>100 kW) and do not adequately address the design considerations/requirements required for scale of very high power solar arrays, such as extremely high stowed packaging efficiency, minimization of mechanical/deployment complexity, high deployed stiffness and strength, and ability to be functionally validated/tested on earth under 1-G. Under funding through the DARPA FAST program, and as presented during the 2009 and 2010 Space Power Workshop Conference, the Boeing Company has been developing a solar array called HPSA (High Power Solar Array) for ultra-high power application. The Boeing Company has designed HPSA for accommodation of a lightweight ultra-thin IMM photovoltaic flexible blanket assembly (under the AFRL's IBIS program), and of a venation-blind type reflective concentrator blanket assembly (under the FAST program). Boeing's HPSA solar array is not similar to the proposed Mega-ROSA embodiment. The HPSA design is very complex, has many mechanisms and cables, and requires multiple motors and heavy mechanisms for deployment of both the support structure and wing structures. Unlike the proposed embodiment, the deployable backbone structure for the Boeing HPSA technology is comprised of many non-orthogonal rotatable structural elements with discrete motorized joints and latches. The entire backbone structure is a motor driven deployment that deploys in an unusually non-orthographic kinematic matter. This complex backbone structure used on the Boeing HPSA technology is very different that the deployable backbone approached suggested by the Mega-ROSA embodiments (such as an accordion folded or telescopic boom backbone structure architectures as presented in the ensuing section).
The only other extremely large area solar arrays were for the International Space Station which were developed and build by Lockheed Martin. Unlike the proposed Mega-ROSA embodiment, the Space Station solar arrays consist of multiple solar array wings, comprised of a central deployable articulated open-lattice boom structure and side photovoltaic blanket assemblies spanning each side of the boom. These deployment solar array wings were then mounted to a non-deployable space-frame rigid truss structure that occupies a very large stowage volume and is not required to be compacted further for stowage like the proposed Mega-ROSA design. The significance of the Space Station solar arrays are that it deployed very large photovoltaic areas to provide high power level and only that.
As the demand for spacecraft power grows to very high power levels (>100 kW), it is desirable to provide a deployable solar array system that permits straightforward scaling up in size to allow the use of larger deployed solar cell areas. It is also desirable to enhance reliability, while at the same time reducing weight and cost, by reducing the number of different component parts and mechanisms required to achieve deployment and adequate deployed performance. Because mechanical components are subject to failure, and must be rigorously tested as an assembled system to validate their reliability; solar array reliability can be increased significantly, while simultaneously reducing cost and mass, by reducing the amount of mechanical components and mechanisms required to deploy and form the array into a deployed structure. A modular approach, where smaller, simpler, manufacturable and testable, building block elements are combined in a straightforward packaging arrangement to form a much larger, yet inherently simple deployable architecture is ideal to meet the very high-power design requirements, and allow the design to remain practical and viable for actual implementation.
The very high power solar array of this invention has been greatly simplified relative to the state of the art by significantly reducing the complexity and number of mechanical parts, and different unique components required for deployment of the extremely large solar cell-populated areas. The invention utilizes the modular building block elements of multiple, similar and simple roll-out solar array deployable solar array “winglet” modules mounted in a repeating fashion onto a central deployable rigid “backbone” structure platform that provides the primary deployed structural stiffness and strength and allows attachment of the solar array to the spacecraft, and stows efficiently into an compact package for launch. A unique aspect of the invention's platform design is its high degree of modularity, scalability and configuration (stowed and deployed form) flexibility.
The roll-out solar array “winglet” module portion of the invention replaces many complex deployable mechanisms required for the unfurling deployment of a typical flex-blanket solar array structure, with a simple ultra-lightweight one-part tubular rolled boom structural element that reliably elastically self-deploys under its own strain energy and is directionally controlled such that it deploys in a known, unidirectional manner without the need for heavy and complex auxiliary actuators to assist deployment or add deploy force. The boom structural element requires no hinges, dampers, complicated synchronization mechanisms, brakes, or motors for deployment, and does not have the parasitic mass associated with the mechanisms typically required by other prior art deployable solar array structures to achieve high deployment force margin. Because the winglet boom structure self-deploys elastically via its own high internal strain energy, it does not require passive (solar) or active (via powered heaters) heating of the boom material to actuate deployment, and provides its own internally-generated high deployment force. The available strain energy for conducting deployment can be maximized to achieve the desired deployment force margin by the use of a highly unidirectional thin fiber-composite layup material for the roll-out boom, because the boom component of this invention is directionally controlled to always unroll in a known and predictable direction, without requiring a special (lower deployment force) bi-stable elastic laminate or elastic memory composite (EMC) material.
The winglet module portion of the invention also enables uniform stowage and secure packaging of the fragile solar cell-populated flexible blanket by maintaining a decoupled arrangement between the blanket longitudinal edges and the deployment structural elements, allowing either a rolled flexible photovoltaic blanket, or an accordion Z-folded flat-package arrangement to be implemented when stowed; and allowing either simultaneous or independent deployment of the boom structure and flexible blanket.
The deployable backbone structure module portion of the invention, provides the mounting interface support and deployment structure for the roll-out winglets; and when fully deployed and latched develops the primary central stiffness and strength for the large very high power solar array wing. In addition to providing the primary structural element, the backbone is capable of packaging extremely efficiently when stowed to enable the spacecraft/launch vehicle integration of such huge power levels. The backbone structure deploys in a controlled, repeatable and synchronized manner, and employs either an articulated or telescopic-type boom configuration; both methods can utilize multiple synchronization/deployment actuation methods (such as spring driven, motorized and cable-pulley).
In more concise terms, a deployable backbone structure for support of one or more pairs of roll-out-solar array winglets is claimed. It comprises a stowage surface from which the backbone structure is deployed. The backbone structure also comprises a deployment boom operable for compact stowage, for extension from the stowage surface via a deployment actuation system, and for providing a stable, secure platform for deployment of one or more pairs of roll-out-solar array winglets. It further comprises a latching system operable for stiffening and stabilizing the deployment boom in its extended position and one or more roll-out-solar array winglets.
The deployable boom may comprise a compactable lattice truss operable for containment within a canister for stowage, for linear extrusion from the containment canister upon extension, and for forming a rigid extended boom upon deployment.
The deployment boom may alternately be comprised of linearly connected backbone beam elements connected via a Z-fold configuration. Each of the hinged beam elements may have a length, first end and a second end. The first end of one of the hinged beam elements may be hingeably connected to the second end of another of the hinged beam elements, thereby forming a hingeably connected linear array of hinged beam elements. Each of the hinged beam elements may also have a synchronized z-fold actuation system operable for effecting the transition between the stowed configuration wherein the hinged beam elements are tightly packed such that their lengths are in immediate proximity to each other, and the deployed configuration wherein the hinged beam elements are rigidly aligned end to end. The hinged beam elements may further comprise a set of planar rectangles operable for stacked arrangement while in the stowed configuration, and rigidly aligned end-to-end in the deployed configuration. Or, the hinged beam elements may further comprise a Z-folded set of nestable deep open section beams, operable for compact nestable arrangement while in the stowed configuration, and for having rigid connection and end-to-end alignment in the deployed configuration.
Alternatively, the backbone beam elements may be telescopically configured with a plurality of telescopic beam elements having a length, first end and a second end. Upon deployment, the first end of one of the each of the telescopic beam elements may be connected to the second end of another of the each of the telescopic beam elements. Each of the telescopic beam elements may also have one or more intermediary structures situated therebetween, operable for sequential extrusion of each of the telescopic beam elements in turn, thereby forming a deployment platform for deployment of the roll-out-solar array winglets. The telescopic beam elements may also have a synchronized telescoping actuation system operable for effecting the transition between the stowed configuration wherein the beam elements are tightly packed such that their lengths are concentric, and the deployed configuration wherein the beam elements are rigidly connected together and aligned end-to end. The telescopic beam elements may further comprise a set of telescopically nestable closed section tubes operable for telescopically nestable arrangement while in the stowed configuration, and for having rigid connection and end-to-end alignment in the deployed configuration.
Other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.